This invention relates to a gas turbine engine and more particularly to the control and modulation of the temperature of the hot gas stream exiting the combustor associated with the engine.
Present day gas turbine engines employed as aircraft power plants operate at high gas temperatures. In fact, one of the key performance factors indicative of the thrust of the engine is combustor exit temperature. To attain and maintain a certain rated thrust, the hot gases exiting the combustor must exhibit a certain average gas temperature level which is typically the highest average gas temperature encountered in the engine. In many instances, this temperature level approaches the temperature limit of the components such as turbine stator vanes, disposed at the combustor exit. Consequently, designers are faced with achieving compatibility between the turbine vanes and the high average temperature of the hot gases exiting the combustor.
Non-uniformity of the temperature of the hot gases in the combustor exit plane is an additional factor which makes the designers dilemma even more acute. The temperature non-uniformities are generally resultant from the geometrical design of the combustor itself. By way of example, the fuel injectors of the combustor contribute to a non-uniform exit temperature distribution in the form of localized temperatures in the plane significantly higher than the average temperature. Specifically, during the combustion process, burning of the air/fuel mixture tends to occur more intensely at the points in the combustor where fuel is injected. Since the airflow through the combustor is at a high velocity, these areas of intense combustion are elongated into hot streaks extending axially along the length of the combustor. In many instances hot streaks may extend axially in the aft direction so far as to encompass turbine stator vanes disposed downstream of the exit of the combustor. Hence, while it is generally correct to say that the designer must design to the average temperature of the hot gases at the exit plane of the combustor, the designer must in fact design the turbine stator vanes to be compatible with the highest single point temperatures of the hot gases at the exit plane. The single point temperatures are hence a significant problem for the designer who has typically responded by applying state of the art cooling techniques. Specifically, the standard approaches have included film cooling of the surfaces of the vane or providing for impingement and internal convection cooling of the vane using compressor discharge air. Use of cooling air in this manner, however, is accompanied by performance decreases in the engine in the form of reduced thrust or greater fuel consumption per unit of thrust output. Furthermore, since the cooling air is introduced in these prior art devices at locations wherein the gases exhibit a high Mach number, mixing losses are high. Additionally, vane designs which utilize cooling techniques, such as impingement, are of complex construction and are high cost components in modern turbine engines.
The degree of non-uniformity of the temperature distribution of the gases at the exit plane of the combustor is highly dependent on the length allowed for combustion. Short combustors tend to produce higher streak temperatures due to inadequate mixing length. Prior art engines hence utilize longer combustors to eliminate the effects of the non-uniformity of the temperature distribution. The present invention is directed at providing a preselected circumferential temperation distribution of the hot gases at the combustor exit plane which reduces the required vane cooling air flows, simplifies the vane mechanical construction, and allows a short combustor design.
Therefore, it is an object of the present invention to achieve compatibility of the turbine stator vanes associated with a gas turbine engine and the hot gases exiting the engine combustor.
It is another object of the present invention to provide for compatibility of the turbine stator vanes and the hot gases exiting the engine combustor without adversely affecting the performance of the engine.
It is still another object of the present invention to eliminate the detrimental effects of a non-uniform temperature distribution of the hot gases exiting the engine combustor upon the turbine stator vanes.
It is yet another object of the present invention to impart a preselected temperature distribution to the hot gases exiting from a foreshortened combustor which distribution is favorable to cooling the turbine stator vanes during engine operation.
Briefly stated, the above and other related objects of the present invention, which will become apparent from the following specification and appended drawings, are accomplished by the present invention which provides, in one form, a gas turbine engine having a hot gas flowing in an annular path partially defined by inner and outer vane shrouds wherein the improvement comprises means disposed upstream of turbine vanes associated with the combustor for establishing a preselected circumferential temperature gradient in the hot gas. The gradient is preselected whereby hot gas at a relatively higher temperature flows through gaps between the vanes and hot gas at a lower temperature flows upon the vanes. The gradient may be established by providing means for admitting air into the combustor of the engine in the form of first and second pluralities of apertures. The first plurality generally provides a locally higher volume of dilution airflow than the second. This first plurality is axially aligned with the turbine vanes, while the second plurality is axially aligned with the aforementioned gaps. Additionally, the first plurality is located axially in a secondary zone of the combustor, but sufficiently upstream of the vanes to induce partial mixing of the dilution air with the hot gases before the partially mixed dilution air impinges upon the vanes. With this axial location the relatively lower gas temperatures of the gradient are present over the entire length of the vanes without using excessive amounts of dilution air. The sinusoidal characteristics of the invention are enhanced by using a number of vanes which are an exact multiple of the number of fuel injectors providing fuel to the engine.